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Publication numberUS7540710 B2
Publication typeGrant
Application numberUS 11/215,392
Publication dateJun 2, 2009
Filing dateAug 30, 2005
Priority dateOct 27, 2003
Fee statusLapsed
Also published asCN1871488A, EP1528343A1, EP1678454A2, US7805945, US8857190, US20060039793, US20070028592, US20100186365, WO2005043058A2, WO2005043058A3
Publication number11215392, 215392, US 7540710 B2, US 7540710B2, US-B2-7540710, US7540710 B2, US7540710B2
InventorsHolger Grote, Heinz-Jürgen Grob
Original AssigneeSiemens Aktiengesellschaft
Export CitationBiBTeX, EndNote, RefMan
External Links: USPTO, USPTO Assignment, Espacenet
Turbine blade for use in a gas turbine
US 7540710 B2
Abstract
A turbine blade or vane for use in a gas turbine is to have as long a service life as possible at high strength. To this end, the turbine blade or vane, according to the invention, has a basic body which is formed from a strengthened cast ceramic material and in which a number of reinforcing elements are placed.
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Claims(8)
1. A turbine blade or vane, comprising:
a basic body formed from a strengthened cast ceramic material and in which a number of reinforcing elements are placed,
wherein the reinforcing element is formed from a ceramic composite material,
wherein the reinforcing element is from a honeycomb-shaped porous material,
wherein the reinforcing element comprises a flat plate arranged in and at a distance from the surface of the basic body, and
wherein the reinforcing element having a plate-shaped design has a number of apertures.
2. The turbine blade or vane as claimed in claim 1, wherein the reinforcing element having an elastic porous structure.
3. The turbine blade or vane as claimed in claim 1, wherein the reinforcing element has a number of beads and thickened portions.
4. The turbine blade or vane as claimed in claim 1, wherein the reinforcing element has a number of beads or thickened portions.
5. The turbine blade or vane as claimed in claim 1, wherein the reinforcing element has a lattice structure.
6. The turbine blade or vane as claimed in claim 1, wherein the reinforcing element has a cross shape, the ends being positioned in the basic body.
7. A gas turbine, comprising:
a turbine blade or vane having a basic body formed from a strengthened cast ceramic material and in which a number of reinforcing elements are placed,
wherein the reinforcing element is formed from a ceramic composite material,
wherein the reinforcing element is from a honeycomb-shaped porous material,
wherein the reinforcing element has a rod shape and extends along a peripheral edge of the basic body.
8. A turbine blade or vane, comprising:
a basic body formed from a strengthened cast ceramic material and in which a number of reinforcing elements are placed,
wherein the reinforcing element is formed from a ceramic composite, honeycomb-shaped porous material,
wherein the reinforcing element has an annular closed shape that extends along a periphery of the basic body.
Description
CROSS REFERENCE TO RELATED APPLICATION

This application is a Continuation in Part of International Application No. PCT/EP2004/012142, filed Oct. 27, 2004 and claims the benefit thereof. The International Application claims the benefits of European Patent application No. 03024560 EP filed Oct. 27, 2003, both of the applications are incorporated by reference herein in their entirety.

FIELD OF THE INVENTION

The invention relates to a turbine blade or vane, in particular for use in a combustion turbine.

BACKGROUND OF THE INVENTION

A combustion space subjected to high thermal and/or thermomechanical loading, such as, for example, a kiln, a hot-gas duct or a combustion chamber of a gas turbine, in which combustion space a hot medium is generated and/or directed, is provided with an appropriate lining for protection from excessively high thermal stressing. The lining normally consists of heat-resistant material and protects a wall of the combustion space from direct contact with the hot medium and from the high thermal loading associated therewith.

U.S. Pat. No. 4,840,131 relates to the fastening of ceramic lining elements to a wall of a kiln. There is a rail system here which is fastened to the wall. The lining elements have a rectangular shape with a planar surface and are made of heat-insulating, refractory, ceramic fiber material.

U.S. Pat. No. 4,835,831 likewise deals with the application of a refractory lining to a wall of a kiln, in particular to a vertically arranged wall. A layer consisting of glass, ceramic or mineral fibers is applied to the metallic wall of the kiln. This layer is fastened to the wall by metallic clips or by adhesive. A wire netting having honeycomb meshes is applied to this layer. The mesh netting likewise serves to prevent the layer of ceramic fibers from falling down. A uniformly closed surface of refractory material is additionally applied by being fastened by means of a bolt. The method described largely avoids a situation in which refractory particles striking during the spraying are thrown back, as would be the case when directly spraying the refractory particles onto the metallic wall.

A ceramic lining of the walls of combustion spaces subjected to high thermal stress, for example of gas turbine combustion chambers, is described in EP 0 724 116 A2. The lining consists of wall elements of structural ceramic with high temperature stability, such as, for example, silicon carbide (SiC) or silicon nitride (Si3N4). The wall elements are mechanically fastened elastically to a metallic supporting structure (wall) of the combustion chamber by means of a central fastening bolt. A thick thermal insulating layer is provided between the wall element and the wall of the combustion chamber, so that the wall element is at an appropriate distance from the wall of the combustion chamber. The insulating layer, which is approximately three times as thick as the wall element, is made of ceramic fiber material which is prefabricated in blocks. The dimensions and the external form of the wall elements can be adapted to the geometry of the space to be lined.

Another type of lining of a combustion space subjected to high thermal loading is specified in EP 0 419 487 B1. The lining consists of heat shield elements which are mechanically mounted on a metallic wall of the combustion space. The heat shield elements touch the metallic wall directly. In order to avoid excessive heating of the wall, e.g. as a result of direct heat transfer from the heat shield element or due to the ingress of hot medium into the gaps formed by the heat shield elements adjacent to one another, cooling or sealing air is admitted to the space formed by the wall of the combustion space and the heat shield element. The sealing air prevents hot medium from penetrating as far as the wall and at the same time cools the wall and the heat shield element.

WO 99/47874 relates to a wall element for a combustion space and to a combustion space of a gas turbine. Specified in this case is a wall segment for a combustion space to which a hot fluid, e.g. a hot gas, can be admitted, this wall segment having a mechanical supporting structure and a heat shield element fastened to the mechanical supporting structure. Fitted in between the metallic supporting structure and the heat shield element is a deformable separating layer which is intended to absorb and compensate for possible relative movements of the heat shield element and the supporting structure. Such relative movements can be caused, for example, in the combustion chamber of a gas turbine, in particular an annular combustion chamber, by different thermal expansion behavior of the materials used and by pulsations in the combustion space, which may arise during irregular combustion for generating the hot working medium. At the same time, the separating layer causes the relatively inelastic heat shield element to rest more fully over its entire surface on the separating layer and the metallic supporting structure, since the heat shield element penetrates partly into the separating layer. The separating layer can thus compensate for unevenness at the supporting structure and/or the heat shield element, which unevenness is related to production and may lead locally to unfavorable concentrated introduction of force.

In particular in the case of walls of high-temperature gas reactors, such as, for example, of gas-turbine combustion chambers operated under pressure, their supporting structures must be protected against a hot gas attack by means of suitable combustion chamber linings. Compared with metallic materials, ceramic materials are ideally suitable for this purpose on account of their high thermal stability, corrosion resistance and low thermal conductivity.

On account of material-specific thermal expansion properties under temperature differences typically occurring in the course of operation (ambient temperature during stoppage, maximum temperature at full load), the thermal mobility of ceramic heat shields as a result of temperature-dependent expansion must be ensured, so that no thermal stresses which destroy components occur due to restriction of expansion. This can be achieved by the wall to be protected from hot gas attack being lined by a multiplicity of ceramic heat shields limited in their size, e.g. heat shield elements made of an engineering ceramic. As already discussed in connection with EP 0 419 487 B1, appropriate expansion gaps must be provided between the individual ceramic heat shield elements, which expansion gaps, for safety reasons, must also be designed so that they are never completely closed in the hot state. In this case, it has to be ensured that the hot gas does not excessively heat the supporting wall structure via the expansion gaps. The simplest and safest way of avoiding this in a gas-turbine combustion chamber is the flushing of the expansion gaps with air, what is referred to as “sealing-air cooling”. The air which is required anyway for cooling the retaining elements for the ceramic heat shields can be used for this purpose.

SUMMARY OF THE INVENTION

The object of the invention is to specify turbine blade or vane which has especially long service life at high strength. Furthermore, an especially low-maintenance turbine blade or vane and a gas turbine having such a turbine blade or vane are to be specified.

With regard to the turbine blade or vane, this object is achieved according to the invention with a basic body which is formed from a strengthened cast ceramic material and in which a number of reinforcing elements are placed.

In this case, the invention is based on the idea that a turbine blade or vane designed for especially long service life should be especially adapted to the external conditions of use. In order to make this possible and provide an especially high number of degrees of freedom for individual adaptation measures, the hitherto conventional production of turbine blades or vanes by pressing is dispensed with and production by casting is now provided instead. However, in a cast ceramic turbine blade or vane, on account of only comparatively low tensile strength in particular in the longitudinal and transverse directions of the turbine blade or vane, the service life of the turbine blade or vane could be limited. In order to therefore enable a turbine blade or vane based on a cast basic body to be used in a turbine for utilizing the structural degrees of freedom achievable with said turbine blade or vane, special measures with regard to the structural reinforcement of the basic body should be taken for long service life and increased passive safety, these measures also increasing the cohesion of the basic body in the event of possible crack formation.

In particular for increased tensile strength and for reducing crack lengths which could occur due to thermal and thermomechanical loads, reinforcing elements are therefore provided which are integrated in the basic body of the turbine blade or vane. In this case, these reinforcing elements should be firmly connected to the turbine blade or vane in order to transfer the material property of the tensile strength of the reinforcing element to the turbine blade or vane. This function is performed by the reinforcing elements positioned inside the turbine blade or vane, these reinforcing elements being integrally cast in the basic body by the ceramic casting material and being firmly connected to the basic body or to the ceramic as a result.

The structural degrees of freedom accompanying the use of a casting technique are advantageously used in the fashioning of the turbine blade or vanes in particular for ensuring, by suitable geometries or local variations in characteristic material properties, an especially high loading capacity even during fluctuating thermal loads on the turbine blade or vanes.

So that a reinforcing element is adapted to the high temperatures to which a turbine blade or vane is exposed, and in addition firmly combines with the ceramic casting material during the casting process, the respective reinforcing element is advantageously formed from a ceramic material, preferably from an oxide-ceramic material having an Al2O3 proportion of at least 60% by weight and having an SiO2 proportion of at most 20% by weight. This material has comparatively high tensile strength and firmly combines with the ceramic casting material on account of the similar mechanical materials during the solidifying. In addition, the thermal expansion of the reinforcing material is similar to the remaining ceramic material of the turbine blade or vane, so that no unfavorable stresses occur in the turbine blade or vane during temperature variations. Furthermore, the reinforcing element may expediently be produced from ceramic fibers such as, for example, CMC materials or from structural ceramic material having a pore proportion of at most 10%.

The reinforcing element can be made out of a ceramic material, with is know for cast filters keeping out slag (waste product) from a cast. This material usually filters due to its porous structure the slag away from the cast. In this utilisation now the porous structure is able used act as a sponge. Ceramic casting material forming the shape of the aerofoil surrounds and flew into the reinforcing element before being hardened. This allows a comparable good bond of the ceramic casting material with the reinforcement element. Similarly effects can be accomplished be having a honeycomb-shaped porous material or bone-structure porous material for the reinforcement element.

The respective reinforcing element is preferably designed like an elongated round ceramic rod in the manner of armoring. In order to integrate a reinforcing element especially firmly in a turbine blade or vane and in order to design the reinforcing element to be as stiff as possible, the latter expediently has beads and thickened portions. The reinforcing element is anchored in the surrounding ceramic material via said beads and thickened portions, as a result of which the tensile strength of the reinforcing elements is transferred to the entire turbine blade or vane. In a rod-shaped configuration, the reinforcing element may in particular have thickened portions at its end region, so that a bone shape is obtained. A positive-locking connection between reinforcing element and basic body is ensured by ends thickened in this way or also by rib-like thickened portions. Alternatively or additionally, this connection may also be made with a frictional grip, for example via a sintering operation or via granulation.

In order to reinforce a turbine blade or vane over the entire surface, a reinforcing element may also expediently be designed in a plate shape, in which case in particular a flat plate arranged in parallel and at a distance from the surface of the basic body may be provided. Here, a plate may be positioned in each case on the side facing the working medium, while a plate for reinforcement is likewise assigned to the cooler side of the turbine blade or vane.

In order to achieve as firm a material bond as possible between a reinforcing element designed as a plate and the surrounding ceramic material, such a plate advantageously has a number of apertures. As a result, the ceramic casting compound can pass into the apertures and also solidify there during the casting process of the turbine blade or vane. In this case, the plate may be designed in particular as a perforated plate, the number, size and positioning of the holes expediently being selected as a function of intended use and material parameters.

In an alternative or additional advantageous embodiment, a reinforcing element of a turbine blade or vane preferably has a lattice structure. In this case, the lattice elements may form a lattice structured with rhombic or square apertures. A reinforcing element may also be formed by a plate which has circular apertures which are positioned at uniform distances apart, so that a lattice-shaped structure is produced.

In order to strengthen or reinforce a turbine blade or vane especially at the sides, a reinforcing element is expediently of rod-shaped design and positioned along a peripheral edge of the turbine blade or vane.

In order to ensure the structural integrity of the turbine blade or vane over its entire periphery even during incipient crack formation, a reinforcing element preferably has a closed annular shape and runs along the periphery of the turbine blade or vane.

In order to increase even further the strength of such an annular reinforcing element and thus also that of the turbine blade or vane and in order to design said reinforcing element and turbine blade or vane in such a way that they are as torsionally rigid as possible, a reinforcing element is expediently designed as a circular ring.

For stabilizing and strengthening the airfoil of a turbine blade or vane, the reinforcing element advantageously has a cross shape, the ends being positioned in the region of the corners of the turbine blade or vane. For suitable bracing of the cross-shaped reinforcing element in the turbine blade or vanes, this bracing increasing the tensile strength, the ends of the cross-shaped reinforcing element may be thickened, so that the reinforcing element is anchored in the turbine blade or vane.

The advantages achieved with the invention consist in particular in the possibility, with recourse to a casting process with the structural degrees of freedom possible as a result, of producing turbine blade or vanes which have especially high tensile strength. By the integration of reinforcing elements in turbine blade or vanes which are made of a cast ceramic material, it is possible to transfer the material properties of the reinforcing elements, such as in particular the tensile strength, to a turbine blade or vane. In this case, the shaping of a turbine blade or vane can be kept flexible. A further advantage consists in the fact that the possibility of selecting various embodiments of reinforcing elements and their positioning in the turbine blade or vane permits individual adaptation to the thermal and mechanical loads acting on a turbine blade or vane. On account of the increased strength of the turbine blade or vanes, the service life of a turbine blade or vane is also prolonged, since the spread of cracks is reduced and the structural integrity of the component (passive safety) is increased.

The advantage of a casting operation consists in the possibility of producing more complex shapes of turbine blade or vanes. Thus, on the one hand, the external basic shape can be varied comparatively easily and at a low cost.

BRIEF DESCRIPTION OF THE DRAWINGS

An exemplary embodiment of the invention is explained in more detail with reference to the drawing, in which:

FIG. 1 shows a half section through a gas turbine,

FIGS. 2 a and 2 b show an exemplary turbine blade and turbine vane of the gas turbine according to FIG. 1,

FIGS. 3 a and 3 b show a turbine blade or vane with plate-shaped reinforcing elements,

FIGS. 3 c and 3 d show a cross section of a turbine profile the surface structure of the reinforcement element,

FIGS. 4 a and 4 b show a turbine blade or vane with a lattice-shaped reinforcing element, and

FIG. 5 shows a turbine blade or vane with a cross-shaped reinforcing element.

DETAILED DESCRIPTION OF THE INVENTION

The same parts are provided with the same designations in all the figures.

The gas turbine 1 according to FIG. 1 has a compressor 2 for combustion air, a combustion chamber 4 and a turbine 6 for driving the compressor 2 and a generator (not shown) or a driven machine. To this end, the turbine 6 and the compressor 2 are arranged on a common shaft 8, which is also referred to as turbine rotor and to which the generator or the driven machine is also connected and which is rotatably mounted about its center axis 9. The combustion chamber 4, designed like an annular combustion chamber, is fitted with a number of burners 10 for burning a liquid or gaseous fuel.

The turbine 6 has a number of rotatable moving blades 12 connected to the turbine shaft 8. The moving blades 12 are arranged in a ring shape on the turbine shaft 8 and thus form a number of moving blade rows. Furthermore, the turbine 6 comprises a number of fixed guide blades 14, which are likewise fastened in a ring shape to an inner casing 16 of the turbine 6 while forming guide blade rows. In this case, the moving blades 12 serve to drive the turbine shaft 8 by impulse transmission from the working medium M flowing through the turbine 6. The guide blades 14, on the other hand, serve to direct the flow of the working medium M between in each case two moving blade rows or moving blade rings following one another as viewed in the direction of flow of the working medium M. A successive pair consisting of a ring of guide blades 14 or a guide blade row and of a ring of moving blades 12 or a moving blade row is in this case referred to as turbine stage.

Each guide blade 14 has a platform 18 which is referred to as blade root and is arranged as a wall element for fixing the respective guide blade 14 on the inner casing 16 of the turbine 6. In this case, the platform 18 is a component which is subjected to comparatively high thermal loading and forms the outer boundary of a hot-gas duct for the working medium M flowing through the turbine 6. Each moving blade 12 is fastened to the turbine shaft 8 in a similar manner via a platform 20 referred to as blade root.

A guide ring 21 is in each case arranged on the inner casing 16 of the turbine 6 between the platforms 18, arranged at a distance from one another, of the guide blades 14 of two adjacent guide blade rows. Here, the outer surface of each guide ring 21 is likewise exposed to the hot working medium M flowing through the turbine 6 and is kept at a radial distance from the outer end 22 of the moving blade 12 lying opposite it by means of a gap. In this case, the guide rings 21 arranged between adjacent guide blade rows serve in particular as cover elements which protect the inner wall 16 or other built-in casing components from thermal overstressing by the hot working medium M flowing through the turbine 6.

In the exemplary embodiment, as shown in FIG. 2, the turbine blade 12 and the turbine vane 14 are configured in a circumferential ring, in which a plurality of turbine blades 12 are arranged in the circumferential direction around the turbine shaft and a plurality of turbine vanes 14 are arranged in the circumferential direction on the inner casing 16.

The turbine blade 12 or vanes 14 are designed in particular for a long service life, so that as little damage as possible occurs due to the external effects, such as the high temperature and flow induced vibrations of the working medium M. To this end, said turbine blade 12 or vanes 14 consist of a basic body 26 which is formed from a cast ceramic material and in which reinforcing elements 30 are integrated. For suitable thermal stability of the reinforcing elements, they are made of a ceramic material or a composite material. To this end, the reinforcing elements 30 can be designed for the effects acting on the turbine blade 12 or vane 14. Various embodiments of turbine blade 12 or vanes 14 with reinforcing elements 30 are presented in FIGS. 3 to 5.

A turbine blade 12 or vane 14 with plate-shaped reinforcing elements 30 is shown in FIG. 3, a reinforcing element 30 being provided in each case for the surface facing the working medium M and the surface facing the cooled side. It can be seen in FIG. 3 that the plate-shaped reinforcing elements 30, for a better bond with a surrounding ceramic, may be provided with a lattice-shaped structure or may be designed as a lattice, in particular as a cross lattice (FIG. 3 a) or as a perforated lattice (FIG. 3 b). The basic body 26 formed as a turbine aerofoil can also created by the porous reinforcement element 30 bond to surrounding ceramic 28 properly because of its own surface structure, independently if it is bone-structured, porous and/or honeycomb-shaped. Even more the surrounding ceramic 28 can flow into the also elastic reinforcement element 30 because of its porous surface structure. Because of its elastic nature the basic body is able to absorb the mechanical tensions occurring during the operation of a gas turbine equipped with such an turbine blade or vane. The porous surface structure of the material known from cast filters is shown in FIG. 3 d.

For especially pronounced reinforcement of the marginal regions of a turbine blade 12 or vane 14, rod-shaped reinforcing elements 30 may be used, as shown in FIG. 4, these rod-shaped reinforcing elements 30 running along the side edges of a turbine blade or vane 26 and being provided with beads or thickened portions (FIG. 4 a) or thickened ends (FIG. 4 b) in order to ensure firm anchoring in the surrounding ceramic 28. In the turbine blade 12 or vane 14 shown in FIG. 5, a cross-shaped reinforcing element 30 is provided in order to brace the structure of a turbine blade 12 or vane 14 in a stabilizing manner, this cross-shaped reinforcing element 30 having thickened portions at each of its ends for anchoring in the ceramic material 26.

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Referenced by
Citing PatentFiling datePublication dateApplicantTitle
US8465276 *Sep 23, 2005Jun 18, 2013Siemens AktiengesellschaftBurner for fluid fuels and method for operating such a burner
US20090061365 *Sep 23, 2005Mar 5, 2009Bernd PradeBurner for fluid fuels and method for operating such a burner
WO2014105108A1 *Mar 15, 2013Jul 3, 2014United Technologies CorporationGas turbine engine component having vascular engineered lattice structure
Classifications
U.S. Classification415/200, 416/224, 416/241.00B, 416/229.00A
International ClassificationF27D1/08, C21B7/06, F27D1/00, F01D5/28, F27D1/10, F27D1/04, F23R3/00
Cooperative ClassificationY10T428/2949, Y10T428/2973, F23R3/007, F27D1/04, F27D1/0033, F27D1/08, F27D1/10, C21B7/06
European ClassificationF27D1/08, C21B7/06, F27D1/10, F23R3/00K, F27D1/00A4, F27D1/04
Legal Events
DateCodeEventDescription
Nov 7, 2005ASAssignment
Owner name: SIEMENS AKTIENGESELLSCHAFT, GERMANY
Free format text: ASSIGNMENT OF ASSIGNORS INTEREST;ASSIGNORS:GROTE, HOLGER;GROB, HEINZ-JURGEN;REEL/FRAME:016984/0743;SIGNING DATES FROM 20050829 TO 20051017
Jan 14, 2013REMIMaintenance fee reminder mailed
Jun 2, 2013LAPSLapse for failure to pay maintenance fees
Jul 23, 2013FPExpired due to failure to pay maintenance fee
Effective date: 20130602